Gas turbine system with displacement compressor of the multiple outer rotor type



Ap 1954 G. K. w. BOESTAD ET AL 2,674,089

GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR TYPE Filed Oct. 29. 1947 7 Sheets-Sheet l A1311] 6, 1954 G. K. w. BOESTAD ET AL 2,674,089

GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULT PE Filed Oct. 29. 1947 IPLE OUTER ROTOR vTY 7 Sheets-Sheet 2 llllllllllll April 1954 G. K. w. BOESTAD ET AL GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR TYPE Filed Oct. 29, 1947 7 Sheets-Sheet 5 April 6, 1954 G. K. w. BOESTAD ET AL 2,674,089

GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR TYPE 1947 7 Sheets-Sheet 4 Filed Oct. 29,

April 6, 1954 e. K. w. BOESTAD ET AL 2,674,089

GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR TYPE 7 Sheets-Sheet 5 Filed Oct. 29, 1947 [419 lik p i 1954 G. K. w. BOESTAD ET AL 2,674,089

GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR TYPE 7 Sheets-Sheet 6 Filed Oct. 29, 1947 INV TORS /z, m; 7%,

)54 AITORNEK April 1954 G. K. w. BOESTAD ET AL 2,674,089

GAS TURBINE SYSTEM WITH DISPLACEMENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR YPE 7 Sheets-Sheet 7 Filed Oct. 29. 1947 u/a/m' 7 7 Patented Apr. 6, 1954 UNITED STATES PATENT OFFICE GAS TURBINE SYSTEM WITH DISPLACE- MENT COMPRESSOR OF THE MULTIPLE OUTER ROTOR TYPE and Percy H. Batten, Racine, Wis., trustees Application October 29, 1947, Serial No. 782,732

Claims priority, application Sweden November 2, 1946 17 Claims. 1' The present invention relates to gas turbine systems and the object of the invention is to provide an improved form of such turbine systems having a low fuel consumption, low weight power ratio as well as small overall dimensions, thus enabling advantageous use of the same not only in stationary and marine power plants but also in prime movers for ail-crafts.

A gas turbine plant contains, as is well known, a compressor means driven by the turbine means and delivering compressed air required as motive fluid in the turbines. In gas turbine systems hitherto used among other compressors also the so-called rotary displacement compressor of sists of or contains a displacement compressor of multiple outer rotor type with at least two outer rotors intermeshing with one common central rotor, said compressor being so arranged as to feed one or several turbines of the turbine means with compressed air.

When constructing the compressor with one common central rotor and two or more outer rotors symmetrically spaced over the periphery of the central rotor, i. e. with the same mutual angle displacement, the Weight ofv the construction for the same delivery capacity will become considerably lower in comparison with the normal two rotor compressor referred to in the prescrew wheel type has been used, such compressor amble of this description. Corr spo comprising two intermeshing rotors mounted in dividing of the turbine means into several units a suitable casing. This two-rotor displacement each acting upon one outer rotor of the comcompressor has the advantage of relatively fiat pressor-possibly over a synchronising gear upon efiiciency curve with simultaneously high peak two or more outer rotors-brings about the posefiiciency over a wide speed range. The comzliloility tthati fi also :hg turbine means may be givelrt l ressor bein of the, ositive type works without e mos e cien imensions and a ow weig Eo-called sur ging phe nomena and with difierent at the same time as the gearing losses are compressure ratios independent of the rotor speed. pleDtely 1or a; least siubstantielillly eliminated.

A disadvanta e of the displacement compressor ue o e ivi ing of e compressor means referred to abo e consists therein that the rotary as well as of the turbine means according to this speed is relatively low within the flat range of invention a gas turbine system is obtained with the efficiency curve and hence the compressor itsmall overall dimensions and above all with a self as well as a turbine directly connected to small frontalarea and lowweight in relation to the the same will 0 erate at low rotary speed, if the output, especially in comparison with previous good efficienoy :i the compressor is to be utilized. embodiments of gas turbine plants with two rotor As it is desirable to give the disks and blades of compressors. the turbine ideal dimensions resulting in a mini- In the following the invention will be more fully mum of weight of the assembly, it will be necesdescribed with reference to the accompanying sary to place a gearing with a suitable reduction drawings illustrating y y Of ple some ratio between the compressor and the turbine. preferred embodiments.

Thereby, however, the reduced weight of the tur- Fig. 1 (shown on two sheets) shows a more or bine Will beco cou t c d by e Weight of less diagrammatic central longitudinal section the added gearing, an d $imu1tane1lS1Y the gear of a gas turbine system with a multiple outer roing will cause an add1t1onal power loss, amounttor compressor and producing power primarily for mg to 13% the transmltted P The propeller propulsion with assisting jet propulpower adsorption of the compressotrn a; gas Sion,

turbine system amounting to abou 3 o e o a turbine output, the power losses in the gearing EE f g i fl sectlon taken on the hne will reduce the net output by from 2-s%.

The present invention consists in such an Fig. 3 1s atransverse section taken on the line rangement of gas turbine plants where the above IIITHI 9 1; mentioned disadvantages are substantially obvi- 4 15 a transverse Sectlon taken on the 1me ated. For this purpose the arrangement accord- IVIV of ing to the invention is substantially character- Fig. 5 is a transverse section taken on the line ized by the fact that the compressor means con- V-V of Fig. l

6 is a transverse section taken on the line VZ-VI of Fig. 1;

Fig. 7 is a schematic longitudinal view of a somewhat modified embodiment of the system as shown in Fig. l, in which the apparatus has a common gas outlet and the separate turbines with appertaining combustion chambers and regeneratcrs are disconnectable, the figure also indicating the path of flow of the motive fluid through the system;

Fig. 8 shows another embodiment in which the shaft of one of the turbines passes freely through the corresponding outer rotor of the compressor so that the complete output of this turbine is transmitted to the output shaft; and

Fig. 9 s s still another embodiment without regenerat which the output of the turbines is balanced to operate the compressor only, the rest of heat energy being utilized for jet propulsion.

In the drawings the reference numeral H3 indicates a compressor of the screw wheel type having intermeshing helical lands and grooves, of which type the compressor shown in Patent No. 2,243,874 granted to Alf Lyshclm is an example. The compressor illustrated is built up of two outer rotors l2 and i4 provided with helical lands and grooves intermeshing in cooperative relation with a common central rotor 15. These three rotors it, it and is are housed in a suitable casing it provided with end walls and 22 at the inlet and outlet ends respectively of the compressor. As will be seen from Fig. 3 the inlet end wall of the compressor is provided with inlet ports 24 and 25 through which air is admitted to the compression spaces of the compressor from the annular air intake channel 28, formed between an outer shell 3i) and inner shell 32, this latter also housing a gearing structure to be described below. The

channel 23 is provided with a screen 34 for sepa-.

rating from the air such particles in the air flow which might have a deteriorating effect on the functioning of the system.

The outer compressor rotors l2 and it are driven each by a separate turbine indicated genorally at it and t2 respectively. The turbine is a full admission multiple stage reaction turbine having a rotor 44 carrying, in the embodiment illustrated, seven rows of rotor blades 46 coacting with the same number of stationary guide or reaction blade rings 43 carried by a turbine casing to. The rotor 54 is mounted coaxially with the outer rotor 2'53 of the compressor and is directly connected thereto by means of the drive shaft 52. Similarly the turbine 42 being of the same type as the turbine 4i], is directly connected to the outer rotor M by means or" a drive shaft 54.

It is characteristic of the type of compressor illustrated that the rotors are operated at high rctative speed with mechanical clearance between the rotors and between the rotors and the casing. In order to maintain the desired clearance between the rotors they are kept in phase relation by means of timing gears 55, 5S and (it, which also perform the additional function of transmitting the excess or net useful power developed by the turbines to a common output shaft. The gear wheels 55 and 53 are secured respectively to the forward or inlet ends of the rotor shaft of the outer rotors i2 and Id. These gear wheels mesh with a central gear til mounted at the forward end of the rotor shaft of the central rotor 16 of the compressor. As will be observed from the drawing the gear as has a larger pitch diameter than the gears 58 and 58. The reason for this is (i ii that in the compressor illustrated the central rotor l6 has more lands and grooves than the outer rotors l2 and I4 and consequently must rotate at a lower speed. In devices of this kind at least the major portions of the helical lands of the central rotor lie within the pitch circle, whereas at least the major portions of the helical lands of the outer rotors lie outside their pitch circles.

In front of gear 60 the shaft of the central rotor carries a sun gear 62 of a planetary gearing comprising planets 64 with which the sun gear 82 meshes and in turn meshing with the stationary outer or ring gear 66 carried by a suitable Web forming a part of the stationary gear case 68. The planets 64 are rotatively mounted on a power output shaft member It! the rearward end of which provides a planet carrier and the forward end of which is connected by any suitable means (not shown) to a variable pitch propeller assembly of any suitable known variety indicated generally at 12.

Advantageously for starting purposes a small single stage turbine 14 may be used, carried in a l separate compartment within the gear case 68 and connected by means of a reduction gear 76 and shaft 18 to the rotor 12 of the outer compressor. This starting turbine is advantageously operated by means of combustion of fuel or by gasifying hydrogen peroxide with permanganate of calcium, these operating agents being carried in a suitable pressure container 80 diagrammatically indicated in the figure.

A separate combustion chamber is provided for heating the motive fluid supplied to each of the turbines. These combustion chambers indicated at 82 and 84 are identical and each comprises an annular body symmetrically surrounding the respective drive shafts 52 and 54. Chamber 82 has at its forward end liquid fuel supplying nozzles indicated at 86 to which fuel is supplied in desired quantity by any suitable form of fuel feeding system, not shown. At its rearward end the chamber 82 delivers heated motive fluid directly to the inlet of turbine 40 through the annular discharge passage 88. The combustion chamber 84 is provided with corresponding details.

The thermal efficiency of a gas turbine system of the kind under consideration is dependent among other things upon the amount of regeneration employed, that is, the amount of heat in the turbine exhaust gases that is transferred to the motive fluid before the latter enters the turbines and thus returned to the cycle. Whether regeneration is to be employed and the amount of regeneration, if employed, depends upon the desired characteristics of the power plant.

The system shown in Figures 1-6 is intended for delivery of power primarily in mechanical form and in order to provide high thermal efliciency of the cycle, for which reason part of the thermal effect in the output gases should be returned to the motive fluid.

For this purpose the arrangement according to the invention is completed with one or more heat exchangers I42, mounted in the outlet channels from the turbines.

The heat exchanger I42 comprises a system of suitably flattened tubes 98 and Hill bent in U form, which tubes can be provided with flanges on the inside as well as on the outside in order to improve the heat transfer. At the inlet ends the tubes are connected to the compressor outlets 36 and 38 through ducts I08 and H0 and headers i081; and mm respectively, whereas the outlet ends of the tubes are connected with the com-' bustion chambers 82 and 84 respectively through the headers II2a and H411 and ducts H2 and I I4 respectively, the latter ducts being primarily connected to the spaces H6 and H8 respectively formed between the outer shells of the cylindrical double-mantled combustion chambers, the air being thereafter conducted to the interior of the combustion chambers through one or more inlet ports I20.

According to a preferred embodiment shown in Figs. 5 and 6 the ducts I08, H0, H2 and I I4 to and from the heat exchanger are placed in the angular spaces between two adjacent turbines and combustion chambers. This arrangement is suitable not only with respect to space limitations but can simultaneously involve a reduction of weight as the outer shells of turbines and combustion chambers form walls of the ducts to and from the heat exchanger so that it is only necessary to add the outer plates I48 (Fig. 5) and the partition walls I48 of the ducts in order to complete the same. The pressures on both sides of the partition walls being substantially the same, said walls I48 might be constructed of thin plate.

As will be observed from Figs. 1-9 of the drawings this entire arrangement lends itself to the production of an elongated assembly of generally oval section and comparatively small frontal area and relatively small vertical dimension so that the assembly is readily adapted to incorporation in an efficient form of nacelle or in the case of large air craft for incorporation within the relatively thick wing of such craft.

The path of flow of the motive fluid through a system as described above is illustrated in Fig. 7 where the dotted arrows show the path of flow from the compressor inlet to the combustion chambers and the solid line arrows show the path of flow from the combustion chambers to the outlet orifices.

The operation of the system assuming it to have been started is as follows:

Air is inducted to the compressor inlet ports 24 through the channel 28 which preferably as shown faces forwardly so as to make available the ramming effect due to the forward flight of the aircraft. The air inducted is carried to the compression chamber created between the casing I 8 and the lands of the outer rotor I2 which intermesh corresponding grooves of the central rotor I6 where the air is compressed and delivered through the discharge port 38. Similarly air inducted from the channel 28 through the inlet port 26 is compressed by the rotors l4 and I8 and delivered through the outer port 38.

From the outlet ports 38 and 38 of the compressor the air flows rearwardly through ducts I08 and H0 to the headers I08u. and 0a and from there through the tubes 98 and I00 in heat exchange relation with exhaust gases to the headers H211 and I Met respectively, and finally through the ducts H2 and H4 respectively to the combustion chambers 82 and 84 through the spaces H8 and H8 into the interior of the combustion chambers through the inlet ports I20. Fuel oil is introduced into the combustion cham her in a suitable quantity through fuel nozzles 88 so that the gases generated by the combustion to the fuel will obtain the inlet temperature required for the turbines 40 and 42. From the outlets 88 of the combustion chambers the gas will flow under expansion through the blade system of the turbines and is then collected in an outer annular channel 90 situated between the heat exchanger and an outer casing 92 extending rearwardly from the part of the construction enclosing the turbines. From the annular channel 90 the gases flow in an inward and backward direction through channels 94 and 96 formed between the flattened tubes 98 and I00 of the heat exchanger to the central discharge space I02 terminating in outlets I04 and ms.

The gases exhausted from the turbines might be employed for producing a jet propulsion effect upon their exit from the rearwardly directed outlets I04 and I06. The extent of this jet propulsion effect is, of course, dependent upon the heat drop of the gases in the turbines and the present system permits the excess energy of the gas after delivery of the power absorbed by the compressor to be utilized in the form of mechanical output power to the output shaft F0 or for jet propulsion by altering the heat drop in the turbines. The greater the heat drop through the turbines the greater the mechanical power delivered by them and the greater the net useful power over and above that required to drive the compressors. Conversely, the less the heat drop through the turbines the less is the available mechanical power and the greater is the residual energy in the exhaust gases available for jet thrust.

While in the embodiment illustrated the compressor shown is of a three-rotor-type in which each of the two outer rotors is driven by a sep arate turbine it is evident that greater capacity can be achieved by using a larger number of similar outer rotors cooperating with the central rotor each of the outer rotors being driven by a separate directly connected turbine. Even in this last mentioned case it is, of course, possible to operate the compressor by means of only some of the turbines or even by means of a single turbine the output of the other turbines being completely delivered to the output power shaft.

Fig. 7 illustrates a modified embodiment of the invention with one common outlet orifice for the turbines. The disengagement of one or the other of the turbines is possible by means of the free wheels I30 and I32 or the like in combination with valves 34 in the gas ducts. This disengagement can be operated manually or automatically and becomes necessary in case of damages in the turbine, combustion chamber or regenerator or if the working condition of the system is such that a gain in thermal efficiency could be obtained thereby.

Fig. 8 illustrates an embodiment in which the compressor is driven by one turbine 40 directly connected to the outer rotor I2 whereas the shaft I40 of the other turbine 42 passes freely through the other outer rotor l4 and delivers useful power. The combustion chambers 82 and 04 respectively can be replaced by a greater number of combustion chambers for each turbine sym metrically arranged around the turbine shafts 52 and 54 respectively.

Fig. 9 illustrates a system designed to produce net useful power entirely by jet thrust. The compressor is of the type illustrated in Fig. 1. Each of the outer rotors is driven by a separate directly connected turbine. The combustion chambers are as shown located between the compressor and turbines enclosing the turbine shafts.

In this embodiment only such a quantity of mechanical energy is required from the motive fluid as is necessary for the compressor and, consequently, the turbines could be constructed with less stages than in the embodiment according to Fig. 1. Furthermore, because of the fact that maximum heat energy in the gases over and above that required to operate the compressor is desira-ble for jet propulsion purposes regeneration is not employed and the air delivered from the compressor is discharged directly to the casings surrounding the combustion chambers, from which the air flows into the combustion chambers to be heated by combustion with the fuel supplied thereto. After having only enough energy extracted therefrom in the turbines to operate the compressor, the gases fiow from the exhausts of the turbines to nozzle-like jet outlets H54 and 165 rearwardly directed of the unit. With such an outlet it is desirable to control the discharge rate and velocity of the gases, and to this end a control member 252 is provided which is axially movable by means of any suitable mechanism such for example as a rack and pinion to vary the area of the jet orifice. By varying this area the back pressure on the turbine can be adjusted so as to likewise vary the pressure and heat drop through the turbine to adjust the amount or" mechanical power developed by it, in accordance with the momentary needs of the system for driving a compressor.

What is claimed is:

1. A gas turbine system comprising turbine means and compressor means driven by said turbine means to supply motive fluid for operating the system, said compressor means comprising a positive displacement rotary compressor having a central rotor and a plurality of parallel outer rotors intermeshing therewith and said compressor having a separate outlet for the hold compressed between said central rotor and each of the outer rotors respectively, said turbine means comprising a plurality of separate turbines corresponding in number to the number of said outer rotors, and separate conduit means including a combustion chamber for connecting each of said outlets to a difierent one of said turbines.

2. A system as defined in claim 1 in which the driving connection between any turbine and its associated rotor includes releasable clutch means, and in which means are provided for separately controlling the flow of motive fluid to said turbines.

3. A system as defined in claim 2 in which said clutch means comprises an overrunning clutch for automatically disconnecting from its associ 'ated rotor any turbine to which the supply of motive fluid is cut ofi.

l. A system as defined in claim 1 in which said turbines are axially spaced from their associated compressor rotors and in which a separate com bustion chamber is located between each turbine and its associated rotor.

5. A system as defined in claim 4 in which said combustion chambers are annular and said turbines include shafts extending through the cores of the combustion chambers.

L. A system as defined in claim 1 in which at least one of said turbines includes a shaft connected to a power absorbing element other than said compressor means.

7. A system as defined in claim. 6 in which one of said outer rotors is hollow and the shaft of said one of said turbines extends through the hollow rotor for connection with another element of the system.

8. A gas turbine system comprising turbine means and compressor means driven by said turbine means to supply a gaseous constituent of motive fluid for operating the system, means communicating with said turbine means for heating said gaseous constituent to provide said motive fiuid, said compressor means including a rotary positive displacement compressor having at least two outer rotors cooperatively meshing with a common central rotor, said turbine means including separate turbines aligned respectively with said outer rotors, regenerator means located on the side of said turbines remote from said co m pressor means, and conduit means for conveying said gaseous constituent from said compressor means through said regenerator means to said heating means.

9. A system as defined in claim 8 in which said conduit means includes duct means located in the space between difierent ones of said turbines for delivering compressed fluid from said compressor means to said regenerator means.

10. A system as defined in claim 8 in which said regenerator is sub-divided into different sections each extending generally in line with one of said turbines, and separate ducts are provided for conducting separate supplies of compressed fluid from said compressor means to the different sections.

11. A gas turbine system comprising a positive displacement rotary compressor having a central rotor and outer rotors parallel and intermeshing with said central rotor at opposite sides thereof, a turbine located in alignment with each oi said outer rotors and axially spaced therefrom, at least one of said turbines being connected with its aligned rotor to drive said compressor, a regenerator located on the side of said turbines remote from said compressor and in the path of the exhaust gases from the turbine, combustion chamber means located in the space between said rotors and said turbines and communicating with the latter, and duct means for conducting compressed fluid delivered by said compressor past said combustion chamber means and turbines to said regenerator and from the latter past said turbines to said combustion chamber means.

12. A system as defined in claim 11 in which said duct means comprises ducts extending through the space between said turbines.

13. A system as defined in claim 11 in which said regenerator means comprises separate sections, one for each of said turbines, and said duct means comprising separate sets of ducts for delivering compressed fluid to different ones of said sections and from the latter to different combustion chamber means connected to different ones of said turbines.

14. A system as defined in claim 13 including valve means for controlling the flow of fluid through said separate sets of ducts.

15. A system as defined in claim 14 in which said compressor means comprises different outlets each in communication with a different one of said outer rotors and in which said separate sets of ducts communicate respectively with different ones of said outlets.

16. A system as defined in claim 15 in which a valved cross connection is provided between said sets of ducts to enable fiuid discharged from a plurality of said outlets to be selectively delivered to one or another of said sets of ducts.

17. A gas turbine system comprising a positive displacement rotary compressor having a central rotor and at least two outer rotors parallel and intermeshing therewith and located at opposite sides of the central rotor, said compressor further having a separate outlet for compressed fluid associated with each of said outer rotors, a separate turbine associated with each of said outer rotors, at least one of said turbines being connected to its associated rotor to drive the compressor, separate combustion chamber means for supplying motive fluid independently to the difierent turbines, and separate duct systems for supplying compressed fluid from said different compressor outlets to the separate combustion chamber means respectively.

References Cited in the file 0! this patent UNITED STATES PATENTS Number Name Date Belluzzo Mar. 9, 1937 Lysholm t Apr. 22, 1941 Lysholm 1 Apr. 28, 1942 Adams May 23, 1944 Gottlieb Aug. 1, 1944 Allen et a1 Oct. 15, 1946 Smith Apr. 15, 1947 Nilsson Sept. 13, 1949 Bonvillian Mar. 21, 1950 Johansson Apr. 10, 1951 

